High Lift Devices . ..... Moreover, engines have a lower efficiency and power-to-weight ratio when ... Meets applicable certification rules in FAA 14 CFR Part 23.
2017 – 2018 Undergraduate Team Aircraft Design Competition Hybrid-Electric General Aviation Aircraft (HEGAA)
Presented by Federal University of Uberlândia Department of Aerospace Engineering Minas Gerais, Brazil
2017 – 2018 Undergraduate Team Aircraft Design Competition Hybrid-Electric General Aviation Aircraft (HEGAA) Presented by Federal University of Uberlândia Department of Aerospace Engineering Minas Gerais, Brazil Team Members
AIAA Numbers
Alexandre Acerra Gil
922668
Eduardo Pavoni Gamba
922678
Gabriel Araújo Hernández
922675
Guilherme Miquelim Caires
922581
Higor Luis Silva
922459
João Paulo Vieira
922563
Kimberlly Costa Carvalho
922564
Luiz Gustavo Santiago
922576
Pedro Brito
922566
Roberto Martins de Castro Neto
922571
Advisor: Dr. Thiago A. M. Guimarães - 498079
Signatures
Table of Contents
1.
Introduction .......................................................................................................................................................... 6
2.
Design Requirements and Proposals .................................................................................................................... 7
3.
Market Research ................................................................................................................................................... 9
4.
Conceptual Design .............................................................................................................................................. 10 4.1. Initial Design Estimate .................................................................................................................................. 11 4.2. Hybrid-Electric Aircraft Design .................................................................................................................... 17 4.3. Design Assumptions ..................................................................................................................................... 19 4.3.1. Typical Aircraft Mission Profile .............................................................................................................. 20 4.3.2. Battery Design ......................................................................................................................................... 21 4.3.2. Internal Combustion Engine (ICE) .......................................................................................................... 22 4.4. Genetic Algorithm Implementation .............................................................................................................. 23 4.5. Aircraft of Comparable Role and Configuration ........................................................................................... 30
5.
Aerodynamic Analysis ....................................................................................................................................... 32 5.1. Conceptual Aerodynamic Design ................................................................................................................. 32 5.1.1. Wing Planform ........................................................................................................................................ 32 5.1.2.
Airfoil Section ...................................................................................................................................... 34
5.1.3.
Optimization of the Wing Planform ..................................................................................................... 36
5.2. High Lift Devices .......................................................................................................................................... 37 5.3. Fuselage ........................................................................................................................................................ 37 5.4. CFD Analysis ................................................................................................................................................ 38 5.4.1. CFD Optimization ................................................................................................................................... 38 5.5. Interference Drag – Wing Fuselage .............................................................................................................. 39 5.6. Winglets ........................................................................................................................................................ 40 5.7. Final Drag Polar ............................................................................................................................................ 41 6.
Performance Analysis ................................................................................................................................... 42 3
6.1. Take-off analysis ........................................................................................................................................... 42 6.2. Aircraft climb ................................................................................................................................................ 43 6.3. Cruise condition ............................................................................................................................................ 45 6.4. Landing ......................................................................................................................................................... 48 6.5. Payload-Range Diagram ............................................................................................................................... 48 6.6. Emergency situation ...................................................................................................................................... 49 7.
Structural Analysis ............................................................................................................................................. 50 7.1. Structural Wing Design ................................................................................................................................. 56 7.2. Structural Empennage Design ....................................................................................................................... 61 7.2.1. Structural Horizontal Tail Design ............................................................................................................ 62 7.2.2.
Structural Vertical Tail Design ............................................................................................................. 64
7.5. Final Material Composition .......................................................................................................................... 69 8.
Aeroelastic Analysis ........................................................................................................................................... 69
9.
Modal Analysis of the Aircraft ........................................................................................................................... 73
10. Stability Analysis................................................................................................................................................ 73 10.1. Longitudinal Static Stability ......................................................................................................................... 73 10.2. Lateral-Directional Static Stability ............................................................................................................... 75 10.3. Longitudinal Dynamic Stability .................................................................................................................... 76 10.4. Lateral-Directional Dynamic Stability .......................................................................................................... 77 11. Center of Gravity ................................................................................................................................................ 78 12. Subsystem Selections ......................................................................................................................................... 80 12.1. Electric .......................................................................................................................................................... 80 12.2. Hydraulic ...................................................................................................................................................... 80 12.3. Fuel ............................................................................................................................................................... 81 12.4. Control .......................................................................................................................................................... 82 13. Landing Gear ...................................................................................................................................................... 84 13.1. Main landing gear ......................................................................................................................................... 84 13.2. Nose landing gear ......................................................................................................................................... 85 4
14. Interior Design .................................................................................................................................................... 85 14.1. Thermal & Acoustic Insulation ..................................................................................................................... 87 14.2. Alternative Cabin Options .................................................................................................................................. 88 15. Geometric Configurations .................................................................................................................................. 89 16. Cost Analysis ...................................................................................................................................................... 91 16.1. Development and Production Costs .............................................................................................................. 91 16.2. Operational Cost ........................................................................................................................................... 94 17. Conclusions ........................................................................................................................................................ 96 18. Bibliography ....................................................................................................................................................... 97
5
1.
Introduction The study of fully or partially electric aircraft has been the subject of discussion between researchers and
engineers at universities and the aeronautics industry over the last years. The need to develop ever more efficient and greener aircraft leads to the motivation to expand technologies and move toward previously unfeasible concepts. Currently, most General Aviation aircraft typically use internal combustion engines (ICE) as a power source. These engines burn fossil fuels with high energy densities, making this type of raw material very advantageous for aviation. However, they are highly polluting, since their burning generates the production of carbon dioxide (CO 2), which is the main responsible gas for global warming. The Air Transport Action Group (ATAG) points out that 2% of anthropometric carbon dioxide emissions come from aviation, and this number only tends to increase along with the number of aircraft in operation [1,2]. Under those circumstances, several targets have been defined in Vision 2020 and AGAPE 2020 for the next few years [3]. Besides the problem of emissions and pollution, the amount of fossil fuels available in the world is limited. Even not knowing about its availability and scarcity in the future, the fossil fuel prices themselves tend to increase in the coming decades due to the fast-growing worldwide energy demands and the uncertain political situation in the Middle East [4]. Thereby, all the situations mentioned above lead to consider and rethink in alternative ways of the power supply. Thus, the introduction of electric propulsion systems has been a great option. First of all, batteries can be used as a power source instead of conventional fuels. However, the battery use itself brings challenges such as the weight on board and its specific energy. Regarding the first one, aircraft are very sensitive when it comes to weight because as it increases on board, the available payload diminishes. Regarding the second one, the batteries currently have low specific energy when compared to high specific energy of conventional fuels. However, studies have shown great results for batteries in the future. Lithium-air batteries may show an impressive theoretical specific energy of 11,680 Wh/kg [5]. But in more realistic numbers, Zn-O2 batteries with specific energy of 400 Wh/kg in 2025 are expected [6]. It is also worth remembering the challenge related to the international aviation regulations, which require that the minimum level of safety compared to the batteries be guaranteed. Moreover, engines have a lower efficiency and power-to-weight ratio when compared to electric motors [7]. Hence, hybrid-electric systems are proposed to balance the advantages of both engine and motor systems, improving 6
the performance. Such systems have potential advantages including low fuel costs, lower vibrations, lower pollution, and reduced noise. After all, it is not an easy task to balance all these interests and to develop a wholly or partially electric propulsion system, but it is of great importance and necessity to carry on the study and improvement of physical limitations, fulfilling the requirements and guidelines for the future. Along those lines, when designing an aircraft, it is important to seek the best aerodynamic efficiency, good stability control, lower associated weight, better aeroelastic behavior, and ease of manufacturing. In this work, it is done a conceptual design of two hybrid-electric aircraft whose design requirements were determined through a Request for Proposal (RFP) published by the committees of the annual design competition sponsored by The American Institute of Aeronautics and Astronautics (AIAA). The initial characteristics and geometries of a conventional aircraft were estimated through the study of historical trends. Then, the hybridization characteristics were evaluated using the series architecture, which affects mainly the range equations and the definition of the propulsive system elements that better fit the project, e.g., the internal combustion engine and batteries. To explore the multidisciplinary characteristics of the aircraft development, it was implemented an optimization based on a genetic algorithm fully integrated with an aerodynamic module with performance and stability constraints, searching for the best values of the geometrical design variables.
2.
Design Requirements and Proposals For the year 2017-2018, the AIAA Request for Proposal (RFP) is for the design of two-member Hybrid-
Electric General Aviation Aircraft family, one for 4 passengers and another one for 6 passengers. The year entry-intoservice (YEIS) is 2028 for a 4-seat model with 1000 nmi of range and 2030 for the 6-seat model with 750 nmi of range. The intent is to have energy storage for takeoff, climb, go-around and emergencies via batteries and electric motors with an engine providing additional power and/or direct propulsion. Moreover, the airframe and propulsion system commonality, by weight, between the 4-seat and 6-seat variant should be 75% or greater of the 4-seater’s empty weight. The other requirements are presented in Tables 1, 2 and 3.
7
Table 1: General requirements for both aircraft. Passenger/pilot weight of 190 lb Baggage weight per passenger/passenger of 30 lb and volume of at least 4 cubic feet per passenger Capable of taking off and landing from different runways (dirt, grass, metal mat, gravel, asphalt, and concrete) Takeoff, and landing performance should also be shown at 5,000 ft above mean sea level (ISA + 18 deg F) as well as for grass & concrete fields at sea level (ISA + 18 deg F)
General Requirements
Minimum cruise speed of 174 knots & Target cruise speed: 200 knots or greater Capable of VFR and IFR flight with an autopilot Meets applicable certification rules in FAA 14 CFR Part 23 Use of engine(s) and electric motor(s) that will be in service by 2028 and document battery energy and power density assumptions based on reasonable technology trends Show airframe and propulsion system commonality of at least 75% between the 4-seater and 6-seater by weight Show the emergency range to get to an alternate airport at the maximum feasible weight from an engine failure at 5000 ft AGL (ISA + 18 deg F) with electric power from batteries alone for both the 4- and 6seat variants Provide systems and avionics architecture that could enable autonomous flight; otherwise, provide a market justification for choosing to either provide or omit this capability Meet 14 CFR 23.67 Climb: One engine inoperative requirements with either propulsion type inoperative if it will be treated as a twin-engine airplane
4-Seat Variant
Table 2: Design requirements for the aircraft of 4 passengers. Crew: 1 pilot + 3 passengers 1000 nmi design range mission with IFR reserves Maximum takeoff and landing field lengths of 1,500’ over a 50’ obstacle to a runway with dry pavement (sea level ISA + 18oF day) Initial climb rate at sea level (ISA+ 18oF) at least 1500 fpm with both electric and fossil fuel propulsion operating
6-Seat Variant
Table 3 - Design requirements for the aircraft of 6 passengers. Crew: 1 pilot + 5 passengers 750 nmi design range mission with IFR reserves Maximum takeoff and landing field lengths of 1,800 ft over a 50 ft obstacle to a runway with dry pavement (sea level ISA + 18oF day) Initial climb rate at sea level (ISA + 18 deg F) at least 1300 fpm with both electric and fossil fuel propulsion operating 8
3.
Market Research For everyone that loves airplanes, the biggest dream is to build one that emits noise as close to zero as possible.
In order to do that, the solution that first come up to mind is a full electric power system, where the noise would be reduced, along with the fuel consumption and air pollution. However, the weight of the set of batteries and generator is so high that the aircraft would be stuck at the ground. Therefore, how could be possible to reduce noise, fuel consumption and air pollution on an aircraft? The answer is simple: hybrid propulsion. The hybrid propulsion combines the power of electric and conventional engines, this way “a plane might taxi to the runway on electricity. During high-demand phases of flight such as takeoff and climb, it could supplement that with energy from a turbine generator to provide lots of power without lots of noise. Once at altitude, the plane could loaf along with just the generator driving electric fans, diverting excess power back to recharge onboard batteries. Then, the plane could descend and land again as a quiet electric”[8]. This reality is not as far away as it looks. According to Engadget[9], Airbus, Rolls-Royce and Siemens are partnering on a hybrid-electric aircraft prototype, the E-Fan X, that will prove the mixture of conventional and electric engines will work. The demonstrator will modify a BAe 146 by replacing one of its gas turbine engines with a 2MW electric motor, followed by a second if everything goes smoothly. It is currently slated to fly sometime in 2020. The aircraft project proposed here would compete with the airplanes in the general aviation category, such as Piper PA 44 180, Diamond DA42, Cirrus SR-22, Cessna 172, Mooney Acclaim and Cessna TTX for the 4PAX aircraft (Dolphin 4000) and Beechcraft Baron, Beechcraft Bonanza, Piper Seneca, Cessna 206, Piper Malibu and Piper MATRIX for the 6PAX aircraft (Dolphin 6000). Figure 1 shows the comparison of takeoff distance between the 4 and 6 seats aircraft. As one can see, both aircraft of the spec have shorter takeoff distance than most of the competitors, which turns possible to takeoff from smaller runways. Figure 1 also shows the comparison of cruise speed and takeoff distance between the 4 seats aircraft. As shown, the cruise speed from the spec is higher than most of the competitors, except for the Cessna TTX. However, the takeoff distance is shorter, which is an advantage over Cessna TTX.
9
Figure 1: Passengers count (on the left) and cruise speed (on the right) vs takeoff distance. Figure 2 shows the comparison of range and takeoff distance between the 4 seats aircraft. As one can see, the range proposed by the spec is greater than most of the competitors. Figure 2 also shows the comparison of cruise speed and takeoff distance between the 6 seats aircraft. As presented, the cruise speed proposed by the AIAA design requirements is among the competitors’ average.
Figure 2: Range (on the left) and cruise speed (on the right) vs takeoff distance. 4.
Conceptual Design This work deals with the design of two hybrid-electric aircraft: one for four passengers (4PAX) and the other
for six passengers (6PAX). Both aircraft are expected to have 75% or greater of communality. Therefore, the first aircraft to be designed was the 4PAX one, since the AIAA specifications require shorter field lengths for takeoff and landing, a higher rate of climb and a more significant range. Thus, after having that aircraft in hands, it is easier to adapt it to the 6PAX configuration, changing the empennage, but keeping the same wing. The design of this hybrid-electric aircraft was obtained using the procedures and theories presented by Venson [10-15], and only the conceptual and preliminary design phases were evaluated. 10
4.1.
Initial Design Estimate First of all, in the conceptual design of the aircraft, several trend tables were used to obtain some initial
estimates. But for that, it was necessary to assign some initial configurations to the aircraft. Thus, four main configurations were assumed: twin-engine with T-tail, twin-engine with conventional tail, single-engine with T-tail, and single-engine with conventional tail. The algorithm used to estimate the design of the aforementioned aircraft considers several coefficients and characteristic values of conventional aircraft presented in historical tables present in the literatures [10-15]. Most of them are represented by a function defined by:
f ( a,W0 , c ) = aW0c where
(1)
f ( a,W0 , c ) represents the characteristic to be calculated such as wetted area ( S wet ) , a
that depend on the type of aircraft under analysis, and
and
c are constants
W0 is the gross weight estimate of the aircraft, which is iterated
throughout the project. This process of iteration and updating of the aircraft weight is performed considering the weight variation and performance during the phases of flight which the fuel consumption is higher: cruise and loiter. In [14,15] the weight ratios per phase are introduced, setting a correlation between them during the aircraft mission. The simplified scheme of iteration flow is presented in Figure 3 (first iteration) and Figure 4 (other iterations).
Figure 3: First iteration for aircraft gross weight estimate. Source: Venson [12]. 11
Figure 4: Other iterations for aircraft gross weight estimate. Source: Venson [12]. Hence, the values used as respective inputs for each aircraft in the algorithm are presented in Table 4. It is worth remembering that these values come from historical tables mentioned above. Table 4: Algorithm inputs for each aircraft configuration. Twin-engine aircraft with T-tail
Twin-engine aircraft with conventional tail
Single engine aircraft with T-tail
Single engine aircraft with conventional tail
ARW
7.8
7.8
7.2
7.2
W
0.75
0.75
0.75
0.75
W [°]
22
22
22
22
ew
0.82
0.82
0.83
0.83
CLmax
1.7
1.7
1.7
1.7
ARHT
4.4
4.3
4.1
39
HT
0.5
0.5
0.5
0.5
VHT
0.96
0.84
0.60
0.45
ARVT
1.0
1.4
1.2
1.5
VT
0.4
0.4
0.5
0.5
VVT
0.071
0.050
0.030
0.030
aSwet
0.2933
0.2933
0.6762
0.6762
cSwet
0.5632
0.5632
0.4884
0.4884
aW0
766
0. 766
0.892
0.892 12
cW0
-0.20
-0.020
-0.047
-0.047
aW
S
1.512
1.512
0.408
0.408
cW S
0.664
0.664
0.804
0.804
aT
0.0116
0.0116
0.0116
0.0116
cT
0.4789
0.4789
0.4789
0.4789
SWet SRef
4.0
4.0
4.5
4.5
C fe
0.055
0.055
0.055
0.055
W1 W0
0.984
0.984
0.990
0.990
W2 W1
0.990
0.990
0.992
0.992
W4 W3
0.992
0.992
0.993
0.993
W6 W5
0.992
0.992
0.993
0.993
aDTO
9.68
9.68
8.23
8.23
aDLND
1.463
1.463
1.524
1.524
a fuselage
0.4088
0.4088
0.4088
0.4088
c fuselage
0.3140
0.3140
0.3140
0.3140
xL d
8.3
8.3
8.3
8.3
xN d
1.3
1.3
1.3
1.3
xC d
0.56
0.56
0.56
0.56
xT d
2.3
2.3
2.3
2.3
Rx
0.34
0.34
0.25
0.25
Ry
0.38
0.38
0.38
0.38
Rz
0.39
0.39
0.39
0.39
CLclean
1.6
1.6
1.5
1.5
CLTO
1.9
1.9
1.8
1.8
CLLND
2.3
2.3
2.0
2.0
0.05
0.05
0.07
0.07
0.44
0.44
0.41
0.41
Rudder Tail Ratio
0.38
0.38
0.36
0.36
Main gear at mac
0.52
0.52
0.50
0.50
0.17
0.17
0.13
0.13
0.34
0.34
0.30
0.30
Aileron wing Ratio Elevator Tail Ratio
Nose gear at length Inner engine span
13
Besides the inputs from Table 4, the algorithm needed an engine to be considered during the performance analysis. Since both aircraft were going to be hybrid, the strategy used here was to get started choosing an electric motor instead of a conventional engine. Searching for commercial and available electric motors, the Siemens’ SP260D was chosen due to its application in electric aircraft, such as the Extra 330LE, also developed by Siemens. Moreover, this electric motor provides 260 kW along with a significant efficiency of 95%, which would be sufficient for both aircraft, and weighs only 50 kg, saving a lot of weight. Further details and technical information are available in [16].
Figure 5: Electric motor SP260D (on the left) and Siemens Extra 330LE (on the right). Source: Endless Sphere. Thus, after running the algorithm, it was released the following outputs for each aircraft, shown in Table 5. Table 5: Algorithm outputs for each aircraft configuration. Twin-engine aircraft with T-tail
Twin-engine aircraft with conventional tail
Single engine aircraft with T-tail
Single engine aircraft with conventional tail
MTOW [kg]
1880.10
1880.10
1466.20
1466.20
Fuel Weight [kg]
297.60
297.60
233.07
233.07
Empty Weight [kg]
1183.30
1183.30
833.97
833.97
399.20
399.20
399.20
399.20
487.97
487.19
433.03
433.14
11.83
11.83
10.74
10.74
crW [m]
1.73
1.73
1.70
1.70
cmacW [m]
1.53
1.53
1.50
1.50
ymacw [m]
2.82
2.82
2.56
2.56
Payload Weight [kg] Structural Weight [kg] bW [m]
14
Sw [m2]
17.94
17.94
16.01
16.01
CLcruise
0.26
0.26
0.22
0.22
CD0
0.02
0.02
0.02
0.02
k2
0.05
0.05
0.05
0.05
xw [m]
2.65
2.72
3.21
3.10
Service Ceiling [ft]
12000.00
12000.00
12000.00
12000.00
Mach (cruise)
0.28
0.28
0.28
0.28
Vstall [m/s]
31.96
31.96
29.88
29.88
W1 W0
0.98
0.98
0.99
0.99
W2 W1
0.99
0.99
0.99
0.99
W3 W2
0.89
0.89
0.89
0.89
W4 W3
0.99
0.99
0.99
0.99
W5 W4
0.98
0.98
0.98
0.98
W6 W5
0.99
0.99
0.99
0.99
268.05
268.05
290.45
290.45
747.39
747.39
680.47
680.47
Range [km]
1228.00
1228.00
1094.60
1094.60
ROC [ft/min]
2483.90
2483.90
1377.20
1377.20
l fuselage [m]
8.52
8.52
8.52
8.52
d fuselage [m]
1.70
1.70
1.70
1.70
S HT [m2]
7.71
6.75
3.53
3.53
crHT [m]
1.77
1.67
1.16
1.31
bHT [m]
5.83
5.39
3.80
3.71
cmacHT [m]
1.37
1.30
0.95
1.00
xHT [m]
6.75
6.85
7.36
7.21
SVT [m2]
4.42
3.11
1.51
1.51
crVT [m]
3.00
2.13
1.58
1.39
bVT [m]
2.10
2.09
0.87
0.87
cmacVT [m]
2.23
1.58
1.74
1.76
Takeoff Distance [m] Landing Distance [m]
15
xVT [m]
5.52
6.39
6.94
7.13
xLanding . gear [m]
4.43
4.43
4.26
4.26
yengine [m]
1.45
1.45
0.00
0.00
I xx [kg∙m2]
7602.20
7602.20
2640.20
2640.20
I yy [kg∙m2]
4926.80
4926.80
3842.30
3842.30
I zz [kg∙m2]
393790.00
393790.00
213500.00
213500.00
xCG (empty
3.81
3.82
4.15
4.14
3.67
3.68
3.89
3.88
1062300.00
1061200.00
893590.00
893750.00
35092.00
35092.00
33082.00
33082.00
116.97
116.97
116.74
116.74
aircraft) [m] xCG (loaded aircraft) [m] Sale Price [US$] Operating Costs per Year [US$/year] Operating Costs per Hour [US$/hr]
Thus, gathering the results from Table 5, the following aircraft configurations were obtained and are presented in Figure 6.
(a)
(b)
(c)
(d)
Figure 6: Sketch of the 4 aircraft released by the initial concept design estimate: (a) twin-engine with T-tail, (b) twin-engine with conventional tail, (c) single engine with T-tail, (d) single engine with conventional tail. Analyzing the AIAA design requirements with the results from Table 5, the aircraft that best satisfies the specifications are the single engine aircraft, either with conventional or T-tail. In both cases, the single engine provides 16
sufficient power for the aircraft to fly and reaches the required performance. So, it was easy to choose only one engine. But for the tail, it goes beyond that. The design of an empennage for any aircraft is extremely important, since it affects the aircraft mass and center of gravity, i.e., the static and dynamic stability. Then, it is necessary to considerer the effects of both empennages arrangement. The conventional tail provides appropriate stability and control, and also leads to the most lightweight construction in most cases, so much so that approximately 70% of aircraft are fitted with it [17]. Moreover, for this configuration, the stabilizer trim is relatively less complex and the vertical tail is usually larger. However, engines cannot be coupled to the rear of the aircraft, what is useful for static stability. Furthermore, spin characteristics can be bad in the case of conventional tail due to the blanketing of the vertical tail, in addition to the downwash of the wing being relatively large in the area of the horizontal tail. On the other hand, the T-tail is heavier than the conventional tail because the vertical tail has to support the horizontal tail. However, the T-tail has advantages that partly compensate this important disadvantage (weight). Because of the end plate effect, the vertical tail can be smaller. In addition, the horizontal tail is more effective because it is positioned out of the airflow behind the wing and is subjected to less downwash. Therefore, it can be smaller. For the same reason, the horizontal tail is also subject to less tail buffeting. As the T-tail creates space at aircraft’s rear, there is enough space to fix the engines, improving static stability. Hence, since the aircraft is going to have a lightweight engine at the nose, considering the information above and aiming an easier structural analysis, the conventional empennage was chosen for this aircraft design. Thus, at this point an overall configuration for the aircraft had been defined. However, it was necessary to improve and refine the aircraft adjusting equations and algorithms for it to become a hybrid-electric one. In other words, the performance equations had to be solved using the hybrid-electric architectures, what would imply in new results and configurations. But, first of all, the propulsion system architecture had to be determined.
4.2.
Hybrid-Electric Aircraft Design Several papers indicate the parallel architecture as the most appropriate arrangement for large aircraft due to
the lower weight associated, since the batteries only feed an electric generator, which helps the main shaft of a turbojet engine, for example. However, the aircraft designed here will be small, carrying four and six passengers, what makes 17
the architecture simpler. Therefore, the series-architecture was chosen for the aircraft, since the engine is already an electric-motor, requiring only electric power to move it. Figure 7 depicts the propulsion system architecture proposed.
Figure 7: Series-architecture used in both 4PAX and 6 PAX aircraft. In the analysis of the participation of each of the energy sources (energy internal combustion (ICE) and batteries), it was considered that ICE would be maintained at its optimum rotation speed and would provide a "constant" power during takeoff, climb and cruise phases. In this way, as the aircraft requires more power than ICE provides, the batteries would be responsible for supplying the remainder of the electrical power, ensuring the operation of the electric motor, which is located on the nose of the aircraft. In the loiter and landing phases, the ICE would no longer be used. Thus, the motor is fully powered by the batteries. For the architecture above, all efficiencies were assumed base on literatures and papers, as presented in Table 6. Table 6: Efficiencies for the hybrid-electric propulsive system. Efficiency
bat inv wiring cont eng generator em prop
Value 90%[18] 95%[19] 97%[20] 99%[21] 60%[22] 95%[23] 95%[23] 85%[24] 18
Now it is necessary to come up with the new performance analysis. First, the classic Bréguet’s range is formulated for a conventional fuel powered aircraft at zero wind conditions, and thrust vector parallel to the airspeed vector, which defines the following:
R=
Wstart
W final
where
V CL 1 dW cT CD W
(2)
cT means the specific fuel consumption. The solution to the integral depends on the flying strategy used (e.g., gradual climb at a constant airspeed and
angle of attack) and the models used to represent the propulsion system and the aerodynamic characteristics. Solutions to the most common flying strategies can be found in many standard textbooks on aircraft performance. But, fundamentally, these solutions use the idea that weight changes gradually throughout the flight. Therefore, the range equation cannot be applied to the hybrid aircraft to be designed here, which uses an electric system (batteries) as energy supply. Thus, Voskuijl presents in [25] a new formulation for the range equation relating the energy stored to the consumption of a hybrid aircraft with parallel architecture. Since the architecture of both aircraft under analysis here is going to be in series, the formulation is reevaluated for the scheme available in Figure 7, which results in the following equation:
R=
contem prop c p H fuel S g (1 − S ) + inv wiring generator g
H batttery H fuel CL CD H fuel + (1 − ) H batttery
( H fuel + (1 − ) H batttery ) gEstart + Wempty + W payload H batttery H fuel ln Wempty + W payload 4.3.
(3)
Design Assumptions Before finishing the concept design, some aspects were added and settled for the development of both aircraft
(4PAX and 6PAX), such as: typical aircraft mission profile, the batteries and an internal combustion engine (ICE) for the hybrid system.
19
4.3.1. Typical Aircraft Mission Profile Based on historical data, FAR – Part 23 and AIAA specifications, some mission characteristics were chosen and are presented in Table 7. The flight level FL120 was chosen to avoid any pressurization issues, since the regulation 14 CFR 91 requires supplemental oxygen system for aircraft flying over FL120. Also, the minimum cruise speed required by the AIAA is 174 knots. Thus, the chosen cruise speed for this project was chosen to be 185 knots, since similar aircraft, such as the Diamond DA42, usually fly around this velocity during cruise. Besides, the rate of climb of 1500 and 1300 fpm were kept the same as specified, and the rate of descent were assumed 900 fpm from trends in [26]. The ranges for both aircraft are also fixed by AIAA and the estimate of endurance in cruise is obtained dividing these ranges by the respective cruise speed. Additionally, the total endurance is an estimate of the time to reach the AIAA requirements (climb and cruise), the assumptions (descent) and the FAR Sec. 125 regulations (taxi, takeoff, approach, loiter and landing). Table 7: Chosen mission characteristics for both aircraft design. Mission Characteristics
4PAX
6PAX
Service ceiling [ft]
12000
12000
Cruise speed [knots]
185
185
Rate of climb [ft/min]
1500
1300
Rate of descent [ft/min]
900
900
Range in cruise [km]
1860
1340
Endurance in cruise [hr]
5.43
3.9
Total endurance [hr]
7.22
5.79
Thus, evaluating the time spent during all phases of flight, reaching the parameters in Table 7, typical flight mission for the 4PAX and 6PAX aircraft regarding altitude and flight time are shown in Figures 8 and 9, respectively. The percentage values are related to the total time spent on the mission, which was calculated as it follows. For 4PAX, for example, during taxi and takeoff, it was considered that the aircraft usually spend a time of 15 minutes. Next, climbing from 0 ft to 12000 ft at a rate of 1500 fpm would take 8 minutes to reach that altitude. For cruise phase, and flying at 185 knots, it would take about 5 hours to reach the specified range of 1860 km. Moreover, it was considered a descent and approach in 50 minutes, and 15 minutes for landing. Summing up all those values, the entire mission would take about 400 minutes. Therefore, it was estimated the percentage values of Figures 8 and 8. 20
14000
Cruise 77%
12000
Altitude [ft]
10000 8000
6000
Descent 3%
Climb 2%
Approach & Loiter 10%
4000 Taxi & 2000 Take-off 4% 0
Landing & Taxi 4%
Figure 8: 4PAX aircraft typical flight mission. 14000
Cruise 71%
12000
Altitude [ft]
10000 8000
6000
Climb 2%
Descent 4% Approach & Loiter 13%
4000 Taxi & 2000 Take-off 5% 0
Landing & Taxi 5%
Figure 9: 6PAX aircraft typical flight mission.
4.3.2. Battery Design The choice of batteries to be used in the aircraft hybrid system was based on the estimates proposed by Hepperle [42] and Girishkumar [43] for the technology available in 2025, which are shown in Table 8. The theoretical specific energy values are not truly feasible in actual applications. The amount of energy that the batteries are able to provide are much less than the expected. In other words, there is an efficiency associated to these chemical reactions involved. Thus, when choosing a battery system, it is very important to consider these effects. Table 8: Specific energy density of current and future chemical battery systems [27]. System
Theoretical specific energy density
Actual battery energy density expected in 2025
Li-Ion
390 Wh/kg
250 Wh/kg
Zn-O2
1090 Wh/kg
400-500 Wh/kg
Li-S
2570 Wh/kg
500-1250 Wh/kg
Li-O2
3500 Wh/kg
800-1750 Wh/kg 21
Later, it was performed an assessment of the feasibility of using such batteries in both aircraft in 2028, as required by AIAA specifications. According to Bruce [28] and Ji [29], the researches with Li-S and Li-O2 batteries are in full development and implementation. However, the reliability and safety of these types of batteries are still unknown, which makes them unviable for aerospace applications expected in 2028. On the other hand, Zn-O2 batteries have been developed since the 1960s and already have medical and telecommunication applications [30]. In addition, companies such as Teck Cominco Metals Ltd. ® and Tesla Motors, Inc.® hold patents for the application of this type of battery in the automotive industry [31,32]. Thus, due to its energy density (higher than for Li-Ion), maturity of their behavior and safety, which is crucial for aerospace applications, the type of battery chosen for both aircraft was the Zn-O2. Using a more conservative approach, the energy density that will be considered for this battery will be 400 Wh/kg. 4.3.2. Internal Combustion Engine (ICE) To choose an internal combustion engine (ICE) that better fits the aircraft selected in Section 4.1, it was compared the viability of two turbo-engines. The first one is a 4-cylinder engine with rated power below the power required for cruising (145 kW) and climb (215 kW), while the second one is a 6-cylinder engine with power output greater than those required in both cruise and climb conditions. Table 9 shows the main characteristics of each ICE. Table 9: Main characteristics of Austro AE330 and Lycoming IO-580 engines. Characteristic Number of Cylinders Rated Power [kW] Mass [kg] Dimensions [m] Fuel consumption [kg/hr]
Austro AE330 4 134 186 0.74 x 0.85 x 0.57 31.20
Lycoming IO-580 6 220 201 0.99 x 0.87 x 0.53 47.68
The degree of hybridization was considered so that the ICE used its maximum capacity during taxi & takeoff, climb and cruise phases. For the remaining phases, aiming to reduce noise during approach, loiter and landing, the ICE would stay in idle, and all the power required for that phase would be supplied by the batteries. To estimate the amount of battery required in each phase of flight, it was considered the typical mission from Figure 8. Thus, during takeoff + climb phase, the aircraft would require from the electric motor about 228 kW of power to reach the rate of climb specified. To provide that 228 kW to the electric motor, the ICE Austro AE330 would provide 134 kW, remaining 102 kW to be provided by the batteries, resulting in a degree-of-hybridization (S) of 0.4126, not 22
considering the efficiencies of the architecture. Considering a takeoff + climb of 8 minutes (i.e., 0.13 hr) and an energy density of 400 Wh/kg for the batteries, it would be necessary 33.15 kg of battery for that phase of flight. Since the fuel consumption of the ICE Austro AE330 is 31.20 kg/hr, for the same time spent, it would be necessary 4.06 kg of fuel. Similarly, the other phases of flight were evaluated. Table 10 shows the comparison of the 4-cylinder engine with the proposed hybridization, and Table 11 shows a 6-cylinder engine with a higher power than required and without hybridization. Table 10: 4-cylinder engine with proposed hybridization. Phase of Flight
S (degree-of-hybridization) Time [hr] Battery Weight [kg] Fuel Weight [kg]
Takeoff + Climb
0.4126
0.13
33.15
4.06
Cruise
0.1092
5.52
226.62
172.41
Loiter
1.0000
0.50
46.83
0.00
Descent + Landing
1.0000
0.22
20.81
0.00
TOTAL
-
6.37
327.41
176.47
Table 11: 6-cylinder with higher power required and without hybridization. Phase of Flight
S (degree-of-hybridization) Time [hr] Battery Weight [kg] Fuel Weight [kg]
Takeoff + Climb
0.0000
0.13
0.00
6.32
Cruise
0.0000
5.52
0.00
266.48
Loiter
0.0000
0.50
0.00
23.68
Descent + Landing
0.0000
0.22
0.00
10.52
TOTAL
-
6.37
0.00
307.01
Analyzing costs, the 4-cylinder engine has an estimated cost of US$106.64 in fuel [33] and US$25.84 for the total recharge of the batteries [34], resulting in a total cost of US$132.48 per flight, while the 6-cylinder engine has a total cost of US$185.42 per flight. Therefore, there is a saving of US$52.94 per flight when using the Austro AE330 engine. In addition, the 4cylinder engine is 38 kg lighter and has a 27% lower volume than the 6-cylinder engine, which justifies its application in conjunction with the hybridization of the aircraft.
4.4.
Genetic Algorithm Implementation In Section 4.1, it was estimated an initial definition for the main characteristics of both aircraft. Next, it was
defined the hybrid architecture and its components. Now it comes up with the idea to implement an algorithm of 23
optimization to vary those characteristics and find better configurations for both aircraft, increasing efficiency and saving weight. Thus, after formulating the new range equation for hybrid-electric aircraft, an optimization algorithm was implemented in Matlab to find the best aircraft configuration. The type of optimization implemented was based on the differential evolution algorithm proposed by Rainer Storn [35]. In this method, 300 generations were created, where each generation contained 10 members per population. However, because of the several iteration variables, the genetic algorithm requires too much processing, which generates a huge computational cost, spending many hours to reach the final result. In this case, for example, it took an average of 15 hours to end the optimization, which justifies the low value of 10 members per population. The input variables for the objective function were generated based on the values of geometry, weight and performance from the results presented in Section 4.1. Table 12 shows the maximum and minimum values used in each optimization variable. Table 12: Maximum and minimum values for each optimization variable. Parameter
Max
Min
Takeoff Engine Power [kW]
200.000
260.000
Climb Engine Power [kW]
200.000
260.000
Wing position along the aircraft’s longitudinal axis [m]
3.238
5.397
Wing span [m]
8.033
13.388
Wing root-chord [m]
1.313
2.188
Wing taper ratio
0.525
0.875
Wing swept angle [°]
0.000
10.000
Wing breaking-point [m]
1.339
2.231
Wing dihedral [°]
0.000
5.000
Wing incidence angle [°]
0.000
3.000
Wing twist [°]
0.000
3.000
Wing profile
1.000
5.000
Horizontal tail span [m]
1.134
1.890
Horizontal tail taper ratio
0.525
0.875
Horizontal tail root-chord [m]
0.794
1.324
Horizontal tail swept angle [°]
0.000
5.000
Horizontal tail profile
1.000
3.000
Vertical tail span [m]
1.126
1.876
Vertical tail taper ratio
0.525
0.875 24
Vertical tail root-chord [m]
0.883
1.472
Vertical tail swept angle [°]
0.000
5.000
Vertical tail profile
1.000
2.000
Fuel Weight [kg]
54.348
380.437
Thus, the genetic algorithm used these inputs to find the best configuration option for the aircraft, meeting the AIAA requirements and achieving a lower gross weight as optimization variable. Different airfoil profiles were used for the objective function to be used in the wings and empennage. All of them were also optimization variables: •
Wing: NACA 23010, NACA 23012, NACA 23015, NACA 23016 and NACA 63415;
•
Horizontal Tail: NACA 0012, NACA 0009 and NACA 2412;
•
Vertical Tail: NACA 0012 and NACA 0009.
For the trimming, aerodynamics and dynamics stability analyses, it was necessary to use a program to aid in the process. For this genetic algorithm, the chosen program was the AVL (Athena Vortex Lattice). An example of how the AVL deal with aircraft geometries is presented in Figure 10.
Figure 10: 4PAX aircraft AVL model. Therefore, the AVL provides aerodynamics and stability results to the genetic algorithm. In addition, it is possible to calculate the performance characteristics using the hybridization ratios. So, it compares these values to the specified criteria shown in Table 13. During the optimization, if the aircraft do not meet all the criteria presented, a penalty is added to an optimization variable. Thus, these aircraft with penalties are discarded over time by the objective function. Figure 11 shows the flowchart used in the optimization process described above. The numbers in the process are associated to the variables available in Table 14. 25
Table 13: Specified criteria for genetic algorithm. 4PAX Min Max Geometrical
Performance
Stability
Min
6PAX Max
ARW
-
13
-
13
ARHT
-
7.5
-
7.5
ARVT
1.5
5.5
1.5
5.5
Takeoff Length [m] Landing Length [m] Rate of Climb [ft/min] Angle of Climb [°] Range [km]
5 1852
457.20 457.20 1500 30 -
5 1389
548.64 548.64 1300 30 -
Emax
10
20
10
20
cm
-
-0.40
-
-0.40
Static Margin Trimmed Angle of Attack [°] Elevator Deflection for Trimmed Condition [°]
10 0 -10
40 5 10
10 0 -10
40 5 10
cn
0.05
-
0.05
-
cl
-
-0.04
-
-0.04
Figure 11: Genetic algorithm flowchart.
26
Table 14: Variables used in the genetic algorithm and represented in Figure 11. Variable Description 1 Wing and Empennage Positions 2
Wing, Empennage and Fuselage Dimensions
3
Wing and Empennage Sweep, Dihedral, Incidence, Twist Angles and Profiles
4
Sea Level Condition
5
CG Position
6
Flap and Control Surfaces Dimensions
7
Flap and Control Surfaces Positions
8
MTOW
9
Moment of Inertia
10
Engine Power/Consumption and Fuel Weight
11
Wing and Empennage CL
12
Non-trimmed Drag Polar
13
Wing and Empennage
14 15
dCl aileron and dCl rudder dCn aileron and dCn rudder
16
Cn and Cl
17
dCL elevator and dCm elevator
18
Structural Weight
19
Full Aircraft
20
Empty and MTOW Static Margins
21
Longitudinal Trim Conditions
22
Lateral–Directional Trim Conditions
23
Range
24
Endurance
25
Rate of Climb
26
Takeoff and Landing Length
27
Service Ceiling
28
Sale Price
29
Operating Costs Per Year
30
Operating Costs Per Hour
31
Trimmed Drag Polar
Cm
Cm
Finally, the genetic algorithm released an optimized aircraft configuration, which satisfies all requirements for the 4PAX configuration. However, this aircraft had to be modified in order to better fit the 6PAX arrangement. A length of 0.75 m was added in the passenger cabin to accommodate the two extra passengers, in addition to the increase in payload. Thus, the optimizer was reshaped so that it provides just a new empennage (horizontal and vertical), since 27
the wing was assumed to be the same. Therefore, the final configurations for both aircraft generated by the genetic algorithm are presented in Table 15. The CATIA models for both configurations are shown in Figure 12. Here it is possible to evaluate the commonality of both aircraft. Since they have the same wing, propulsion components (ICE, electric motor, batteries), and part of the fuselage, the commonality of by weight obtained is 96%, which minimizes the development, certification, and manufacturing costs a significant amount.
Figure 12: 4PAX aircraft CATIA model. Table 15: Final optimized configurations for both hybrid aircraft. 4PAX 6PAX MTOW [kg] 1855.10 1966.20 Fuel [kg] 176.57 130.63 Empty [kg] 1279.30 1236.80 Weight Payload [kg] 399.20 598.80 Structural [kg] 474.38 497.50 TZFW [kg] 1713.80 1861.70 Battery [kg] 370.63 309.71 bW [m] 13.38 13.38 [m]
1.31
1.31
1.19
1.19
ARW
12.84
12.84
W
0.70
0.70
W [°]
0.080
0.080
S w [m2]
13.95
13.95
CLmax
1.51
1.51
CD0
0.024
0.024
k2
0.046
0.046
crW
cmacW Wing
[m]
28
xw [m]
2.88
3.52
Y-Position Break [m] W [°]
3.61
3.61
4.70
4.70
iw [°]
2.00
2.00
2.00
2.00
Maximum Celling [m] Vcruise [m/s]
5486.40
5486.40
94.14
94.14
M cruise T /W
0.28
0.28
0.11
0.10
Vstall [m/s]
33.20
34.18
Takeoff Distance [m] Range [km] Landing Distance [m] ROC [ft/min] Takeoff Angle [°] Length [m] Cabin [m] Nose [m] Tail [m] Cabin Diameter [m] S HT [m2]
442.16 1869.80 442.16 1558.50 9.83 8.12 2.40 2.21 3.51 1.70
533.62 1404.00 533.62 1311.90 7.96 8.87 3.15 2.21 3.51 1.70
2.23
1.84
ARHT
6.31
5.36
0.76
0.75
HT
0.57
0.55
VHT
0.65
0.52
bHT [m]
3.75
3.14
0.59
0.59
xHT [m]
7.58
8.11
HT [°]
0.00
0.00
SVT [m ]
0.84
0.63
ARVT
3.27
4.27
0.70
0.54
VT
0.44
0.41
VVT
0.022
0.016
bVT [m]
1.66
1.64
0.50
0.38
7.66
8.32
CLmax
Performance
Fuselage
with take-off flap
crHT Horizontal Tail
[m]
[m]
cmacHT
2
crVT Vertical Tail
cmacVT
[m]
[m]
xVT [m]
29
Control Surfaces
Landing Gear
Electrical Engine
VT [°]
0.00
0.00
Aileron Root Chord [m] Aileron Tail Chord [m] Aileron Span [m] Y-position Aileron [m] Elevator Root Chord [m] Elevator Tail Chord [m] Elevator Span [m] Rudder Root Chord [m] Rudder Tail Chord [m] Rudder Span [m]
0.31 0.29 0.50 5.99 0.19 0.11 3.75 0.35 0.16 1.66
0.31 0.29 0.50 5.99 0.18 0.10 3.14 0.27 0.11 1.64
xLanding . gear − Nose [m]
1.11
1.15
xLanding . gear − Main [m]
4.26
4.43
Takeoff Power [kW] Climb Power [kW] Cruise Power [kW] Descent Power [kW]
240.00 228.12 150.43 24.56
230.00 216.99 151.04 31.98
0.85
0.85
prop
Center of Gravity
Stability
4.5.
Empty
xCG [m]
3.21
3.73
MTOW
xCG [m]
3.35
3.96
TZFW
xCG [m]
3.30
3.92
cn
0.054
0.066
cl
-0.11
-0.11
cm
-0.38
-0.42
Empty Static Margin MTOW Static Margin Trim Angle [°] Elevator Trim Deflection [°]
35.15 16.72 2.58 -1.79
32.63 12.89 2.89 -1.13
Aircraft of Comparable Role and Configuration Searching for other aircraft of the same category, it was found models such as the Cessna TTx, Cirrus SR22
and Diamond DA42. These aircraft present similar characteristic to the hybrid-electric aircraft designed here. As an exemplification, they are compared to the hybrid-electric aircraft of 4PAX in Table 16. In addition, they are placed over each other to show the differences in size, as show in Figure 13.
30
Table 16: Aircraft of comparable role and configuration. Characteristics Diamond DA42 Cirrus SR22 Cessna TTx Wing span [m] 13 11 11 Fuselage length [m] 8.56 8.28 7.68 Height [m] 2.49 2.72 2.74 Range [km] 1693 1289 2352 Endurance [hr] 8.8 4.5 5.25 Cruise speed [km/h] 326 226 435 Takeoff field [m] 776 600 390 Landing field [m] 620 620 805 MTOW [kg] 1999 1111 1633 Passengers 3 3 3 Fuel [kg] 231 220 280 Service ceiling [m] 5500 4100 7620 Climb Rate [ft/min] 1337 720 1400
Hybrid 4PAX 13.38 8.12 2.96 1875 7.22 343 442 442 1855 3 176 3660 1559
Figure 13: Aircraft of comparable role and configure overpainted. Therefore, the aircraft developed in this work meet the AIAA requirements and also is competitive in the aviation market. Since it presents good general and performance characteristics, its main differential certainly would be lower fuel consumption because of the hybrid-electric system on board, attracting new customers. The aircraft presents a longer and larger nose compared to the other aircraft because all the propulsive system, which includes batteries, inverter, ICE, generator, controller and electric motor, was assumed to be placed at the front of the aircraft. Due to its similarity, that “big” nose inspired the name of the aircraft: Dolphin. Thus, the aircraft of 4 passengers was named Dolphin 4000 and the one of 6 passengers, Dolphin 6000. 31
5.
Aerodynamic Analysis The goal of the aerodynamic design was to find the best compromise between drag and lift, especially to the cruise
phase by ensuring a high lift to drag ratio to maximize the range of the aircraft and to reduce the battery’s weight. Some challenges were initially identified by the aerodynamic sector in reason of some conflicting requirements of the spec. The requirements are:
5.1.
•
High minimum climb rate of 1500 ft/min;
•
Same wing geometry to both aircraft to ensure high commonality (imposed by the team);
•
Considerable battery weight to sustain a minimum cruise speed of 174 knots. Conceptual Aerodynamic Design
5.1.1. Wing Planform The first step was the determination of the wing planform to be used in both aircraft, and some initial considerations about the structural twist and aerodynamic tailoring to guarantee good spanwise lift distribution to minimize the induced drag factor (K). Firstly, some aircraft with fairly recent designs were analyzed to study the aspect ratio tendency. Eight aircraft were considered: DA-42, DA-62, Cirrus SR22, Cessna TTx, Bonanza S35, Piper Seneca, Cessna 206 and the Cessna 210. The geometric characteristics of the eight wings are shown in Table 17.
Older Aircraft
Recent Aircraft
Table 17: Wing characteristics of similar aircraft. Aircraft
Year
Wingspan [m]
Wing Area [m2]
Aspect Ratio
DA-62
2012
14.57
17.1
12.41
DA-42
2002
13.56
16.3
11.29
Cirrus SR22
2001
11.68
13.5
10.11
Cessna TTX
2004
11
13.1
9.24
Piper Seneca
1971
11.96
19.3
7.38
Cessna 206
1962
10.92
16.3
7.32
Cessna 210
1957
11.2
16.2
7.73
Bonanza S35
1947
10.21
16.8
6.21
Due to the high climb rate imposed by the spec, a range including high values for aspect ratio, as the more recent aircraft shown in Table 17, was considered. The choice for a high aspect ratio is beneficial due to good glide characteristics, low induced drag factor and high maximum lift coefficient (closer to the airfoil). It was stipulated a
32
range for AR between 12 and 13.5. Although high AR leads to low flutter velocities, this characteristic is not critical to this project since the flight regime is below Mach 0.4. With the goal of ensuring low manufacturing cost, elliptical planforms were eliminated, moreover excluding unfavorable stalling characteristics presented in these types of geometry. Therefore, only aerodynamic surfaces defined by terms of straight lines were considered and three configurations: rectangular tapered, double tapered and triple tapered were pondered. According to Thomas [F. Thomas, 1984] the combination of trapezoidal section together with the appropriate distribution of structural twist can produce a planform that performs nearly as well as an elliptical distribution, overcoming the stalling problems. Thomas states that a double tapered wing with −3° twist in the outer section seems to offer the best compromise, which can be seen in Figure 14. These results were obtained for 𝐴𝑅 = 25 but vary little with aspect ratio. Thus, a structural twist of −3° was assumed.
Figure 14: Effect of wing planform and twist on induced drag [F. Thomas]. Thomas in his book also states that both theoretical investigations and practical experience suggest that a double tapered wing having a taper ratio of 0.4:0.8:1 and a taper break at y/s = 0.6 yields especially good results. Triple tapered wings have featured only designs where there is an intense need for better approximate of an elliptical planform, as can be seen in some sailplanes, such as the Discus, ASH 25 and Nimbus 4. To continue ensuring low manufacturing costs, a double taper configuration was assumed to be enough to ensure good lift distribution. Still analyzing Table 17, it is possible to observe a tendency for wing area in the range of 13 to 17 m2 for more recent aircraft. Taking into consideration that almost all aircraft structure will be manufactured using composite 33
materials, which leads to low empty weight, a decision for a small wing area was taken. A small range of 13.5 to 14 m2 was assumed for the wing area. The maximum wing root chord was also limited to 1.4 m, in order to reduce the percentage of the turbulent boundary layer over the chord of the entire wingspan. This decision was based on light aircraft that have used this strategy to keep as much laminar region as possible to reduce drag. So far, it has been determined the range for the wing’s aspect ratio, the structural twist, the number of sections of the wing, wing area and an initial idea about how the taper ratio and break point can occur. Table 18 shows all the ranges for the wing’s characteristics. Table 18: Wing characteristics and range of variables Characteristic Value Unit
5.1.2.
AR Twist Area
12 - 13.5 3 13.5 - 14
--° m²
Number Sections Max Chord Taper Taper Brake Point
2 1.4 Double tapered 0.6
--m --y/s
Airfoil Section The selection of the airfoils was done considering the following requirements:
•
to exhibit low drag in high-speed flight (low lift coefficient);
•
to provide a high maximum lift coefficient for low landing and takeoff speeds;
•
to contribute to gentle stall characteristics;
•
to be as thick as possible to allow deep spars and high torsional stiffness; The last three requirements do not conflict and are easy to solve by using thicker airfoils, which tend to be
superior in low-speed flight; however, the first requirement is usually solved by adopting thin airfoils and increasing performance at high speeds. Although the four requirements are conflicting, relatively good performance can be reached by balancing out thin and thick airfoils at certain regions of the wing. The choice of airfoils was thought so that flow separation occurs first on the inboard sections of the wing. This prevents premature loss of aileron control and reduces the tendency of the aircraft to fall off on one wing and enter into a spin. To solve this problem, the incidence of the wing’s root airfoil had to be set so that the maximum lift coefficient was exceeded first on the inboard section of the wing. 34
Although it was stated that a taper ratio of 0.4:0.8:1 yields especially good results for lift distribution (elliptical), the CL-distribution of rectangular tapered wing gives the best stalling behavior. Nonetheless, having an elliptical wing geometry makes the 𝐶𝐿𝑚𝑎𝑥 to be reached along the entire wing all at once, and the stall characteristics may be problematic. This way, the final taper ratio of the wing is still open and might be a little different of 0.4:0.8:1. To improve the stall characteristics of the double tapered wing, different airfoils were considered. In this respect, the airfoils, especially in the outer wing, was selected to gentle trailing edge stall behavior. Considering the MTOW of both aircraft, 1855 kg for the Dolphin 4000 and 1966 kg for the Dolphin 6000 (see Table 15 in Section 4.4), and taking into consideration the atmospheric conditions at 12000 ft, the 𝐶𝐿𝑐𝑟𝑢𝑖𝑠𝑒 is around 0.38 for the Dolphin 4000 and 0.40 for the Dolphin 6000 for minimum cruise velocity. Based on the family of airfoils of some similar aircraft and on the considerations made before, airfoils NACA five-digit and 6-series were selected to further study of the aerodynamic characteristics. An especial attention was given to the airfoils of the five-digit family due to the big experimental data available [Abbot & Doenhoff] that match close Reynolds number of the mean aerodynamic chord of the wing at cruise speed. Table 19 shows some characteristics of eight airfoils for two Reynolds numbers, four of them thought to the inner sections of the wing and the last four to the outer sections. Almost all airfoils chosen were designed to have lift coefficient of 0.3. Although the lift coefficient of the aircraft is closer to 0.4 than it is to 0.3, such decisions were based on the fact that at least 10% of the total lift of the aircraft will come from the fuselage for cruise configuration. Table 19: Airfoil’s characteristics. CLmax Re=6e6 1.7
Cm at designed 𝒄𝒍 -0.03
𝑪𝒅
LE radius
0.3
CLmax Re=3e6 1.5
0.0059
2.48%
1.85
-0.01
0.0060
2.82%
N23015
Designed Location Root
N23016
Root
0.3
1.78
N64415
Root
0.4
1.48
1.6
-0.3
0.0050
1.59%
N63415
Root
0.4
1.53
1.63
-0.3
0.0052
1.59%
N23010
Tip
0.3
1.71
1.82
-0.01
0.0054
1.10%
N23012
Tip
0.3
1.6
1.72
-0.04
0.0050
1.58%
N64210
Tip
0.2
1.35
1.41
-0.18
0.0043
0.72%
N64212
Tip
0.2
1.4
1.45
-0.18
0.0045
1.04%
Airfoil
Designed 𝒄𝒍
The analysis of the airfoils from Table 19 reveals that though the NACA 6-series presents smaller values of drag coefficient than the NACA five-digit, due to the laminar bucket, the small leading-edge radius leads to lower maximum lift coefficients in reason of rapid acceleration or deceleration, resulting in sharp pressure gradient [S. 35
Farokhi]. Furthermore, the 6-series airfoils produce considerable aft load if compared to the five-digit family, which requires larger horizontal tails; consequently, more drag can be generated when considering the entire aircraft. Although the drag coefficients are slightly higher for the five-digit, the good surface quality offered by the carbon fiber together with a good manufacture precision can help to reduce the friction drag. Therefore, the airfoils NACA23016 and NACA23015 were considered to the inner sections of the wing and the airfoils N23012 and NACA23010 to the outer sections.
5.1.3.
Optimization of the Wing Planform To determine the final planform of the wing, an advanced differential evolution algorithm was used. In this
process, modules of performance, stability & control and aerodynamics (coupled with the AVL software) using Matlab was employed to help meet all the criteria imposed (climb ratio, range, etc.). The range of all parameters and goals of the optimization process are shown in Table 20. Table 20: Optimization range of all parameters and optimization goals Variable
Lower Limit
Upper Limit
Unit
AR
12
13.5
---
S
13.5
14.5
m²
𝒄𝒓𝒐𝒐𝒕
1.2
1.4
m
Structural Twist
-3
-3
º
Taper
0.4:0.7:1
0.7:0.8:1
---
Taper Break Point
0.45
0.65
---
Variable
Goals
Aerodynamic Eff.
>13.5
---
𝑪𝑳𝒎𝒂𝒙 w/o flaps
minimum 1.7
---
𝑪𝑳𝒄𝒓𝒖𝒊𝒔𝒆
0.38 (D4000) & 0.40(D6000)
---
Cruise Velocity
> 174
Knots
The optimization process was repeated several times to ensure that the same solution was reached at the end of each run. The final geometry was later analyzed using other potential solvers such as the Non-Linear Lifting Line coupled with viscous corrections to check the aerodynamics coefficients obtained using AVL. The final planform of the wing can be seen in Figure 15.
36
Figure 15 - Final wing planform 5.2.
High Lift Devices In order to fulfill the requirements stipulated by performance sector, a flap was designed to reach 𝐶𝐿𝑚𝑎𝑥 of
2.14. The main idea was to introduce the structure only in the first section of the wing to reduce the manufacturing complexity and with 20% of the local chord. The theory of lift induced by partial span flaps below stall presented by Roskam was used to verify the 𝐶𝐿 increase due to the chosen flap. The 𝐶𝐿 increase should be enough to achieve the ratio of climb specified previously and to land at low speeds. ∆𝐶𝐿 was estimated using the Lowry and Polhamus's method and reached a value of 0.47 for a small deflection of 15°, which is high enough to provide the lift requested.
5.3.
Fuselage The design process of the fuselage, as mentioned before, had the goal of producing at least 10% of the total
lift of the aircraft during cruise flight. Besides that, the right position of all internal components had to be ensured to guarantee correct CG location with an external shape that generates minimal drag force. A design close to a slender shape was thought to reduce the drag coefficient by avoiding any detached flow over the entire fuselage and without impairing visual capacity inside the cockpit. The initial design was based (inspired) on the aircraft Pipistrel Panthera. The sketch of the fuselage is shown in Figure 16.
Figure 16: Views of the fuselage. 37
5.4.
CFD Analysis To better analyze the flow characteristics and take into consideration the boundary layer and turbulence
effects, some computation fluid dynamics simulations (CFD) were performed using the complete aircraft geometry. The two-equation model k-epsilon, one of the most traditional turbulence models still in use, was employed in Ansys Fluent® solver. This model was chosen due to a better facility of convergence and lower computational cost inherent to its numerical methodology. Besides that, the k-epsilon turbulent model allows us to work in the log-law region (30
